The assignee of the present invention manufactures and deploys spacecraft for, inter alia, communications and broadcast services. Market demands for such spacecraft have imposed increasingly stringent requirements on spacecraft payload operational capacity, and such requirements may consequently require increased power demands on the spacecraft. For spacecraft that produce power through photovoltaic systems, i.e., solar panel arrays, such increased power demand may be met by increasing the efficiency of the solar cells in the solar panel arrays or increasing the photovoltaic area of the solar panel arrays.
Launch vehicle compatibility is a second requirement faced by a spacecraft designer. The increased performance requirements are only advantageously met if compatibility with conventional, commercially available launch vehicles is maintained. Accordingly, a spacecraft, as configured for launch, is desirably made compatible with the mass properties and fairing envelope constraints of such launch vehicles as, for example, Ariane V, Atlas XEPF, Proton, and Sea Launch. As such, solar panel arrays must be capable of being stowed in a launch configuration that is compatible with the launch vehicle, usually folded flat against a side of the spacecraft, and must then be capable of being deployed into an on-orbit operational configuration after the spacecraft reaches orbit. The major dimensions of a solar panel in such a solar panel array may thus be constrained in size so as to be approximately the same size as the panel of the spacecraft main body to which it attaches, e.g., in typical orbital spacecraft, the “north” and/or “south” panels (the panels of the main body facing the north/south directions, respectively, when the spacecraft is on-orbit). Since such solar panels are typically too small to satisfy the power requirements of many spacecraft, multiple instances of such solar panels are often connected together in a foldable array that is folded against a panel of the spacecraft body and that is configured to be transitioned into a deployed configuration once the spacecraft is on-orbit
FIGS. 1A through 1C depict a style of solar panel array that features a one-dimensional array of solar panels. Each panel in such an array is connected with the adjacent panels by hinge interfaces on opposing sides of the panel. Through rotation of the panels about these hinge interfaces by 180 degrees in alternating directions, the panels may be extended in a concertinaed fashion from a stowed configuration, such as is shown in FIG. 1A, into the deployed configuration, such as is shown in FIG. 1C, along a deployment axis; such an array may be referred to as an in-line solar panel array. FIG. 1B shows the spacecraft of FIG. 1A midway through deployment of the in-line solar panel array.
While such an arrangement may, in theory, be used to supply an in-line solar panel array of any length, too great an increase in length results in an undesirable increase in the rotational moment of inertia of the spacecraft as each additional panel results in additional mass placed at further and further distances from the spacecraft main body.
FIG. 1D shows a schematic of another style of solar panel array that may include two additional solar panels connected to a common in-line panel in an in-line solar panel array. The additional solar panels, or “side” panels, may be connected to the common in-line panel along opposing free edges of the in-line panel using additional hinge interfaces. When fully deployed, such solar panel arrays may have a “T” or cruciform appearance, and may be referred to as cruciform solar arrays.
Cruciform solar panel arrays may feature a side panel restraint mechanism, such as is described in U.S. Pat. No. 6,010,096 to Varouj Baghdasarian, one of the present inventors. U.S. Pat. No. 6,010,096 is hereby incorporated by reference in its entirety. Such side panel restraint mechanisms may include a latch that restrains the side panels from rotational movement until the hinge that includes the side panel restraint mechanism has latched into a locked configuration. For example, a cruciform solar panel array may include a side panel restraint mechanism or mechanisms in the hinge interface or interfaces between the in-line panel to which the side panels are connected and an adjacent in-line panel. The side panel restraint mechanism may prevent substantial rotational movement of the side panels until after the in-line panels have locked into their deployed positions. One of the side panels may also feature a side panel restraint mechanism that prevents the other side panel from releasing until after the first side panel has locked into its deployed position; this prevents intra-side panel collisions. Such arrangements, however, are limited since the side panel restraint mechanisms and the side panels that they restrain must be connected to the same in-line panel. For example, cruciform panel arrangements may generally support only five or six solar panels.
For various reasons, such as design complexity and the likelihood of inter-panel collisions, spacecraft designers utilizing deployable rigid panel solar panel arrays have generally adopted either in-line solar panel arrays or cruciform solar panel arrays for use with spacecraft. With the increasing power needs of spacecraft, however, there is a need for larger sizes of solar panel arrays that in-line and cruciform solar panel arrays may not be capable of providing.